The high temperature of inlet gas stream air entering high pressure turbine nozzles and flowing over outer surfaces of individual vanes of the nozzles in a gas turbine engine has required cooling of the vane airfoil sections in order to maintain vane temperatures within the present material capability. Cooling is commonly provided by forming the vanes as hollow airfoils and providing vent holes from the hollow interior through which a cooling gas, typically air, is forced. The gas desirably forms a film over at least a portion of the airfoil surface and thereby cools or at least insulates such surface. The film cooling injection location is extremely important on the suction side (convex surface) of the airfoil where the hot gas stream can become supersonic. Performance considerations have driven film cooling to be introduced on the airfoil surface at locations where the hot gas stream has a low velocity and near the leading edge of the airfoil section. The selection of cooling film injection locations is a trade-off between performance and cooling of the airfoil. Performance losses are directly proportional to the square of the main stream Mach number at the injection locations. Therefore, the impact on engine performance is significantly different when comparing performance when coolant is injected in a region where the Mach number is about 0.3 as opposed to injection in a region where the Mach number is about 1.0. However, when injection occurs in a low Mach number region, the cooling film may degrade to a point of being ineffective prior to reaching the vane trailing edge. In order to compensate for such degradation, it is necessary to increase the flow of coolant, but such increased flow adversely affects the temperature profile out of the combustor and adversely affects engine performance. Accordingly, coolant injection is often a trade-off of performance against cooling and component life.
With some high curvature airfoil sections, the gas film or vent holes are oriented angularly so as to reduce the gas film injection angle. The reduced angle improves the ability of the film to flow along the airfoil surface. If the film does not flow along the surface, i.e., if it is dissipated in the gas stream, then cooling is ineffective. Film blow-off occurs if the strength of the injected coolant relative to the strength of the gas stream, i.e., the blowing rate, is incorrect for the coolant injection angle. It has also been proposed to turn the cooling gas through a large angle, e.g., between 135 and 165 degrees, using a curved admission tube before injecting the cooling gas at an angle of between about 15 and 45 degrees with respect to the airfoil surface, to try to force the film to remain on the vane surface over greater distances. However, this arrangement has been applied to airfoils having relatively continuously curved suction sides which do not introduce rapid velocity changes. More particularly, this proposed arrangement has been demonstrated to be effective only for blowing rates of between about 0.37 and 0.70. For blowing rates above 0.70, the curved tube was found to be less effective in film cooling than straight tube injection. This above approach is discussed in detail in NASA Technical Paper 1546 published in 1979 and entitled "Influence of Coolant Tube Curvature in Film Cooling Effectiveness as Detected by Infrared Imagery", by Papell, Graham, and Cageao. In general, it is believed that blowing ratios greater than 1.1 are less effective in film cooling.
The development of blunt leading edge airfoils creates more severe film cooling requirements. With such airfoils, a high curvature section exists immediately downstream of the normal film injection point. Conventional injection processes are ineffective to maintain the cooling film on the airfoil surface over such high curvature regions. Furthermore, the velocity of the high temperature gases over high curvature regions approaches supersonic velocities and contributes to the degradation of the cooling film due to large free stream turbulence.